1. Field of the Invention
The present invention relates to a gas turbine blade rotating airfoil and more specifically to a means for relieving stress proximate the blade platform trailing edge.
2. Description of Related Art
In a gas turbine engine, turbine blades are exposed to severe operating conditions and as a result, the blades are susceptible to high cycle fatigue (HCF), low cycle fatigue (LCF), and thermal mechanical fatigue (TMF) cracking in the region where the airfoil meets the blade platform. In order to minimize the exposure of this region to HCF, LCF, or TMF cracking, it is important to isolate this region from the main load path of the airfoil. The cycling can be driven by either temperature or resonance.
As hot combustion gases pass through the turbine section of the engine, blade temperatures can rise well above the operating level of the blade material. In order to compensate for this temperature effect, turbine blades are cooled. Typical cooling configurations have a cooling medium entering the blade through an attachment region and traveling radially outward through the platform to the airfoil. Once in the airfoil, the cooling medium may make several radial passes through the airfoil before exiting through a plurality of holes in either the airfoil surface, blade tip, or blade trailing edge. In order to maximize the amount of gases passing through the turbine and the overall blade weight, the airfoil sections are relatively thin. In contrast, blade platform sections are much thicker and have a higher mass in order to provide adequate support for the airfoil and its associated loads. Therefore, given exposure to a generally uniform combustion gas temperature, the platform region, having a greater mass, is less responsive to thermal changes than the airfoil, creating effectively a thermal fight at their interface, resulting in high thermal stresses.
Normal engine operations can result in cycling of these high thermal stresses, which can lead to crack initiation and potentially damaging crack propagation.
The other principal driver in HCF crack propagation in the region where the airfoil meets the platform is resonance. That is, the airfoil experiences a vibration due to the surrounding turbine and combustion environment. More specifically, this could be due to low order frequency modes, the effects of the quantity of upstream or downstream blades and vanes, or effects from the combustion system.
Manufacturers of prior art turbine blades have attempted to address the thermal stress issues by providing a cutback to the platform, to allow the platform to respond for actively to temperature fluctuations. Two examples of prior art blades contain this cutback, 15 and 46, shown in FIGS. 1 and 2, respectively. The prior art blade in FIG. 1 attempts to address crack propagation by incorporating a cutback along the trailing edge side of the platform. However, this cutback does not extend into the stress field created by the turbine blade airfoil, and therefore cannot redirect the mechanical stresses away from the blade trailing edge while allowing the platform trailing edge region to be more responsive to thermal fluctuations. The prior art blade shown in FIG. 2 also attempts to address the concern of crack propagation by directing the load path of airfoil 40 away from the trailing edge side 48. This is accomplished by configuring cutback 46 such that it is oriented at an angle with respect to the mean camber line of airfoil 40, with cutback 46 beginning on the concave side of the platform and exiting the platform on the trailing edge side. Furthermore, cutback 46 extends to a depth that enters the load path of airfoil 40 to further reduce the vibratory effects of airfoil 40 at the trailing edge region. The preferred embodiment for incorporating this cutback configuration, given its complex geometry, while maintaining structural integrity of the airfoil/platform region during the casting process, would be to machine the cutback into the platform region during blade final machining. However, this machining step requires additional time and machine set-up, and is more costly than if a cutback having a similar effect could be incorporated into the casting or into an existing machining step, where no additional cost is incurred.
Attempting to incorporate this type of cutback into a casting could result in casting flaws and excessive scrap parts since the cutback is only along a portion of the platform, thereby creating a non-uniform section of the blade platform to cool after the blade has been cast.
What is needed is a gas turbine blade having reduced vibratory and thermal stresses at the region between the airfoil trailing edge and adjacent platform, wherein the means for obtaining these reduced stress levels ease blade manufacturing.
In order to solve the problems presented by the prior art, the present invention discloses a turbine blade that has an airfoil to platform interface that is configured to minimize the thermal and vibratory stresses. Therefore, exposure to the conditions that are known to cause high cycle fatigue and low cycle fatigue cracks are minimized. This is accomplished by incorporating a channel in the platform trailing edge that extends from the platform concave face to the platform convex face. Extending the channel across the entire width of the platform removes unnecessary material from the blade platform, which lowers overall blade pull on the turbine disk, resulting in increased life of the blade attachment region. This channel can be incorporated into the turbine blade through either the casting or machining process. The channel, which has a portion having a constant radius, crosses into a line of stress created by the turbine blade airfoil load and redirects the mechanical stresses away from the blade trailing edge while allowing the platform trailing edge region to be more responsive to thermal fluctuations.
It is an object of the present invention is to provide a gas turbine blade with lower thermal and vibratory stresses.
It is another object of the present invention to incorporate a means for lowering the thermal and vibratory stresses while reducing manufacturing complexity.
It is yet another object of the present invention to reduce overall turbine blade weight while increasing blade attachment life.
In accordance with these and other objects, which will become apparent hereinafter, the instant invention will now be described with particular reference to the accompanying drawings.